Wing and method of manufacturing

ABSTRACT

A wing includes a wing box including interconnected spars, an interior system installed within the wing box, an opposed pair of skins fastened to and covering the wing box, wherein one of the skins closes out the wing, and a plurality of fastening systems configured to fasten the skins to the spars and provide protection from electromagnetic effects, wherein each one of the fastening systems includes a threaded fastener, a nut plate including a body and a cover, and a nut enclosed within the nut plate between the body and the cover, wherein the nut is restricted from rotation within the nut plate about a nut plate axis and is free to move linearly within the nut plate orthogonal to the nut plate axis.

PRIORITY

This application is a divisional of U.S. Ser. No. 15/245,769 filed onAug. 24, 2016.

FIELD

The present disclosure is generally related to wings for aircraft and,more particularly, to electromagnetic effect compliant aircraft wingsand methods of manufacturing the same.

BACKGROUND

Composite structures are used in a wide variety of applications,including in the manufacture of airplanes, spacecraft, rotorcraft andother vehicles and structures, due to their high strength-to-weightratios, corrosion resistance and other favorable properties. In theaerospace industry, composite structures are used in increasingquantities, for example, to form the wings, tail sections, fuselage andother components, due to their better specific strength and stiffness,which translates to weight savings, which translates into fuel savingsand lower operating costs.

As an example, composite aircraft wings may utilize upper and lowerouter composite wing skin panels, commonly referred to as “skins,” thatare mechanically attached or bonded to an internal frame. The internalframe may typically include reinforcing structures such as spars, ribsand/or stringers to improve the strength and stability of the skins. Theskins may be attached to the spars, and the spars provide structuralintegrity for the wings. In addition, many aircraft wings may be used asfuel tanks (e.g., a fuel tank is defined inside the wing), which may becontained between front and rear spars.

However, composite structures in aircraft do not readily conduct awaythe extreme electrical currents and electromagnetic forces generated bylightning strikes. Therefore, aircraft with composite structures, suchas composite wings, may be equipped with protection againstelectromagnetic effects (EME) from lighting strikes. For example,conductive media may be provided on a surface to dissipate lightningcurrent away from underlying metal structures and/or fastener systems.In addition, gaps between fastener parts (e.g., two-piece fasteners) andgaps between fastener parts and structural members may be filled withdielectric sealant that provides EME protection. Even if some current isnot diverted, the sealant prevents arcing and sparking across the gaps.

However, current EME protection architectures for composite wings arecomplex and expensive. As an example, the processes of installing thetwo-piece fasteners and applying the sealant requires extensivemanufacturing labor and is performed in confined spaces. For example,the process of manufacturing the wing typically involves match drillingthe spars and the skins, removal of the skins from the spars for surfacefinishing, and realignment of the skins to the spars to close out thewing. Access to the now closed out wing for installation of the fastenerparts, installation of other interior systems and injection of thesealant is gained through access holes formed in the lower outer skin,which is inefficient and potentially dangerous for the laborer.Moreover, the sealant adds weight to the aircraft. While the weightadded to a single fastener system might seem insignificant, applying thesealant to tens of thousands of fasteners in a single aircraft can addhundreds of pounds.

Accordingly, those skilled in the art continue with research anddevelopment efforts in the field of aircraft wings and, in particular,EME compliant wings.

SUMMARY

In one embodiment, the disclosed wing includes a wing box includinginterconnected spars, an interior system installed within the wing box,and an opposed pair of skins fastened to and covering the wing box,wherein one of the skins closes out the wing.

In another embodiment, the disclosed wing includes a wing box includinginterconnected spars or interconnected spars and ribs, an interiorsystem installed within the wing box, an opposed pair of skins fastenedto and covering the wing box, wherein one of the skins closes out thewing, and a plurality of fastening systems configured to fasten theskins to the spars and provide protection from electromagnetic effects,wherein each one of the fastening systems includes a threaded fastener,a nut plate including a body and a cover, and a nut enclosed within thenut plate between the body and the cover, wherein the nut is restrictedfrom rotation within the nut plate about a nut plate axis and is free tomove linearly within the nut plate orthogonal to the nut plate axis.

In another embodiment, the disclosed fastening system, to fasten a skinto a spar of a wing, includes a threaded fastener configured to bereceived through a skin fastener hole in the skin, a nut plateconfigured to be coupled within a spar fastener hole of the spargenerally aligned with the skin fastener hole, wherein the nut platecomprises a body and a cover, and a nut enclosed within the nut platebetween the body and the cover, wherein the nut is restricted fromrotation within the nut plate about a nut plate axis, and the nut isfree to move linearly within the nut plate orthogonal to the nut plateaxis.

In yet another embodiment, the disclosed method for making a wingincludes the steps of: (1) forming a wing box including interconnectedspars, and a plurality of spar fastener holes formed through the spars,each one of the spar fastener holes comprising a spar fastener holediameter, (2) forming skins comprising a plurality of skin fastenerapertures, each one of the skin fastener holes comprising a skinfastener hole diameter, wherein the spar fastener hole diameter islarger than the skin fastener hole diameter, (3) installing an interiorsystem within the wing box, (4) installing nut plates within each of thespar fastener holes, wherein each one of the nut plates includes asleeve configured to be received and retained within an associated oneof the spar fastener holes, a flange extending radially from the sleeveand defining a nut receiving recess, a dome cover extending axially fromthe flange opposite the sleeve and defining an interior chamber and anut at least partially received within the nut receiving recess andenclosed within the cover, wherein the nut is restricted from rotationwithin the nut plate about a nut plate axis and is free to move linearlywithin the nut plate orthogonal to the nut plate axis, (5) sandwichingthe wing box and enclosing the interior system between the skins withthe skin fastener holes generally aligned with the spar fastener holes,wherein a skin fastener hole center axis of each one of the skinfastener holes is not coaxially aligned with a spar fastener hole centeraxis of each one of the spar fastener holes, (6) installing fastenersthrough each one of the skin fastener apertures and the sleeve of eachone of the nut plates, (7) coaxially aligning a nut axis of the nut withthe skin fastener hole center axis, (8) fastening the fasteners to thenut of the nut plates, (9) providing protection from electromagneticeffects, and (10) closing out the wing.

Other embodiments of the disclosed apparatus and method will becomeapparent from the following detailed description, the accompanyingdrawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an aircraft;

FIG. 2 is a schematic block diagram of aircraft production and servicemethodology;

FIG. 3 is a schematic side perspective view of one embodiment of thedisclosed wing;

FIG. 4 is a schematic side elevation view, in section, of one embodimentof the disclosed fastener system;

FIG. 5 is a schematic partial side elevation view, in section, of oneembodiment of the disclosed wing and fastener system;

FIG. 6 is a schematic partial side elevation view, in section, ofanother embodiment of the disclosed wing and fastener system;

FIG. 7 is a schematic enlarged partial side elevation view, in section,of another embodiment of the disclosed wing and fastener system; and

FIG. 8 is a flow diagram of one embodiment of the disclosed method formaking a wing.

DETAILED DESCRIPTION

The following detailed description refers to the accompanying drawings,which illustrate specific embodiments and/or examples described by thedisclosure. Other embodiments and/or examples having differentstructures and operations do not depart from the scope of the presentdisclosure. Like reference numerals may refer to the same feature,element or component in the different drawings.

Illustrative, non-exhaustive embodiments, which may be, but are notnecessarily, claimed, of the subject matter according the presentdisclosure are provided below.

FIG. 1 is a schematic illustration of an exemplary embodiment of anaircraft 1200, such as in the form of an airplane 1216 (e.g., a fixedwing aircraft). As illustrated in FIG. 1, the aircraft 1200 includes twoor more wings 1218. Each wing 1218 may incorporate one or moreembodiments of the disclosed wing 100 (FIG. 3) and disclosed fastenersystem 200 (FIG. 4). The aircraft 1200 also includes a fuselage 1220 anda tail 1222, for example, that includes horizontal stabilizers 1224 anda vertical stabilizer 1226. The wings 1218, horizontal stabilizers 1224and/or vertical stabilizer 1226 may take the form of an airfoil (e.g.,includes an airfoil-shaped body in cross-section). As further shown inFIG. 1, each wing 1218 includes a leading edge 1228, a trailing edge1230, a tip end 1232, a root end 1234 and an internal frame 1236. Eachwing 1218 may also include one or more fuel containment regions, such asa fuel tank 1240.

Embodiments of the wing 100, the fastener system 200 and method 500 formaking the same disclosed herein may be described in the context of anaircraft manufacturing and service method 1100, as shown in FIG. 2, andthe aircraft 1200, as shown in FIG. 2.

FIG. 2 is an illustration of a flow diagram of an exemplary embodimentof the aircraft manufacturing and service method 1100. Duringpre-production, the illustrative method 1100 may include specificationand design, as shown at block 1102, of the aircraft 1200, which mayinclude design of the wing 100, and material procurement, as shown atblock 1104. During production, component and subassembly manufacturing,as shown at block 1106, and system integration, as shown at block 1108,of the aircraft 1200 may take place. Production of the wing 100, asdescribed herein, may be accomplished as a portion of the production,component and subassembly manufacturing step (block 1106) and/or as aportion of the system integration (block 1108). Thereafter, the aircraft1200 may go through certification and delivery, as shown block 1110, tobe placed in service, as shown at block 1112. While in service, theaircraft 1200 may be scheduled for routine maintenance and service, asshown at block 1114. Routine maintenance and service may includemodification, reconfiguration, refurbishment, etc. of one or moresystems of aircraft 1200 (which may also include modification,reconfiguration, refurbishment, and other suitable services).

Each of the processes of illustrative aircraft manufacturing and servicemethod 1100 may be performed or carried out by a system integrator, athird party, and/or an operator (e.g., a customer). For the purposes ofthis description, a system integrator may include, without limitation,any number of aircraft manufacturers and major-system subcontractors; athird party may include, without limitation, any number of vendors,subcontractors, and suppliers; and an operator may be an airline,leasing company, military entity, service organization, and so on.

As shown in FIG. 1, the aircraft 1200 produced by the exemplary aircraftmanufacturing and service method 1100 may include an airframe 1202, aplurality of high-level systems 1204 and an interior 1206. Examples ofthe high-level systems 1204 include one or more of a propulsion system1208, an electrical system 1210, a hydraulic system 1212 and anenvironmental system 1214. Any number of other systems may be included.

Although the aircraft 1200 shown in FIG. 1 is generally representativeof a commercial passenger aircraft having wings 1218 that incorporateone or more embodiments of the disclosed wing 100, the teachings of theembodiments disclosed herein may be applied to other passenger aircraft,cargo aircraft, military aircraft, rotorcraft, and other types ofaircraft or aerial vehicles, as well as aerospace vehicles, satellites,space launch vehicles, rockets, and other aerospace vehicles, as well asautomobiles and other land vehicles, boats and other watercraft,structures such as windmills, or other suitable structures.

Apparatus, systems and methods embodied herein may be employed duringany one or more of the stages of the aircraft manufacturing and servicemethod 1100. For example, components or subassemblies corresponding tocomponent and subassembly manufacturing (block 1106) may be fabricatedor manufactured in a manner similar to components or subassembliesproduced while aircraft 1200 is in service (block 1112). Also, one ormore apparatus embodiments, method embodiments or a combination thereofmay be utilized during production stages such as component andsubassembly manufacturing (block 1106) and system integration (block1108), for example, by substantially expediting assembly of and/orreducing the cost of the aircraft 1200 while complying withelectromagnetic effects (EME) requirements. Similarly, one or moreapparatus embodiments, method embodiments or a combination thereof maybe utilized, for example and without limitation, while aircraft 1200 isin service (block 1112) and during maintenance and service stage (block1114).

FIG. 3 is a schematic illustration of a side perspective view of anexemplary embodiment of the disclosed wing 100, for example, a compositewing, such as in the form of the aircraft wing 1218 (FIG. 1). In theillustrated embodiment, the wing 100 includes one or more spars 102 anda plurality of stiffened outer wing skin panels, generally referred toas skins 130. The wing 100 may also include a plurality of ribs 128.When utilized, the ribs 128 are connected to the spars 102, for example,extending approximately perpendicularly between adjacent pairs of spars102. The spars 102 or the spars 102 and the ribs 128 form an internalframe 134 of the wing, such as the internal frame 1236 (FIG. 1) of theaircraft wing 1218 (FIG. 1).

Each spar 102 includes a first end 104, a longitudinally opposed secondend 106 and an elongated body 108. The body 108 may be continuous (e.g.,unitary) body or segmented. As an example, the illustrated wing 100includes a front spar 102 a and a rear spar 102 b. The front spar 102 ais positioned lengthwise along a leading edge 110 of the wing 100, suchas in the form of the leading edge 1228 of the aircraft wing 1218 (FIG.1). The rear spar 102 b is positioned lengthwise along a trailing edge112 of the wing 100, such as in the form of the trailing edge 1230 ofthe aircraft wing 1218. As another example, the wing 100 may alsoinclude one or more intermediate spars (not explicitly illustrated). Theintermediate spars are positioned lengthwise (e.g., at intermediatelocations) between the front spar 102 a and the rear spar 102 b. Thespars 102 provide strength to the wing 100 and may carry axial forcesand bending moments.

In an exemplary embodiment, each one of the spars 102 may be attached toa fuselage of an aircraft, such as the fuselage 1220 (FIG. 1) of theaircraft 1200 (FIG. 1). As an example, the first end 104 of each of thespars 102 is configured for attachment to the fuselage. In otherembodiments, the spars 102 may be attached to other suitable structuresof the aircraft.

The spars 102 extend from the fuselage in a lengthwise direction from aroot end 114 toward a tip end 116 of the wing 100, such as from the rootend 1234 (FIG. 1) toward the tip end 1232 (FIG. 1) of the aircraft wing1218 (FIG. 1). In the illustrated embodiment, the second end 106 of eachof the spars 102 extends toward the tip end 116 of the wing 100 and/orterminates proximate (e.g., at or near) the tip end 116.

In the illustrated embodiment, the wing 100 includes one or more fuelcontainment regions 118 disposed in the wing 100, such as in the form ofthe fuel containment region 1238 (FIG. 1) of the aircraft wing 1218(FIG. 1). In an exemplary embodiment, the fuel containment region 118includes a fuel tank 120, such as in the form of the fuel tank 1240(FIG. 1). However, in other embodiments, the fuel containment regions118 may include a fuel cell or another suitable fuel containment regionor structure.

In an example, and as shown in FIG. 3, the fuel containment region 118,such as in the form of the fuel tank 120, has fuel containmentboundaries 122 a, 122 b, 122 c, 122 d that form the perimeter of thefuel containment region 118. Although the example fuel containmentregion 118 shown in FIG. 3 has a four-sided, generally rectangularconfiguration, in other examples, the fuel containment region may beformed in other suitable configurations.

In an embodiment of the wing 100, the front spar 102 a and the rear spar102 b are closer to the tip end 116 than intermediate spar, which mayhave a second end that terminates near a middle portion of the fuelcontainment region 118. However, in other embodiments, the second end ofthe intermediate spar may terminate at longer or shorter lengths withinthe fuel containment region 118.

In the illustrated embodiment, the front spar 102 a and the rear spar102 b extend in the lengthwise direction through both a wet section 124of the wing 100, containing the fuel containment region 118, and througha dry section 126 of the wing 100, not containing the fuel containmentregion 118. As used herein, the term wet section means a fuel barrierarea where fuel is contained and the term dry section means an areawhere no fuel is contained.

In an embodiment, portions of one or more of the spars 102 may form astructural wall of at least one of the one or more fuel containmentregions 118. For example, a portion of the front spar 102 a may form thestructural wall of the fuel containment region 118 along the fuelcontainment boundary 122 d. A portion of the rear spar 102 b may formthe structural wall of the fuel containment region 118 along the fuelcontainment boundary 122 b. The portions of the spars 102 forming thestructural wall are interior portions of the spars 102.

In the illustrated embodiment, the plurality of ribs 128 are attachedsubstantially perpendicular to and between the one or more spars 102. Asan example, each one of the plurality of ribs 128 intersects with thespars 102. The plurality of ribs 128 stabilizes and provides support tothe wing 100. In an embodiment, a portion of the plurality of ribs 128separates the one or more fuel containment regions 118 within the wing100.

In the illustrated embodiment, the skins 130 include one or morestiffened upper outer wing skin panels, generally referred to as anupper skin 130 a, and one or more stiffened lower outer wing skinpanels, generally referred to as a lower skin 130 b. In FIG. 3, theupper skin 130 a is depicted as being transparent in order to betterillustrate the internal frame 134 of the wing 100, as shown with brokenlines.

The upper skin 130 a and the lower skin 130 b cover or sandwich the oneor more fuel containment regions 118, the one or more spars 102 and theplurality of ribs 128 between the upper skin 130 a and the lower skin130 b. The plurality of ribs 128 may transfer load among the spars 102and the upper skin 130 a and the lower skin 130 b.

In the illustrated embodiment, the wing 100, such as in the form of theaircraft wing 1218, includes or contains a spar wing box, or simply awing box 132, also referred to as a ladder assembly. The wing box 132includes the internal frame 134 or substructure of the wing 100 andincludes (e.g., is formed by) the interconnected spars 102 and ribs 128.The wing box 132 may include the fuel containment region 118. The upperskin 130 a and the lower skin 130 b cover or sandwich the wing box 132;thus, closing out the wing box 132.

As an example embodiment, the spars 102 (e.g., the front spar 102 a, therear spar 102 b and/or any intermediate spars) may be made (e.g.,formed) of a composite material. As an example, the spars 102 may bemade of fiber-reinforced polymer, or fiber-reinforced plastic, thatincludes a polymer matrix reinforced with fibers, such as carbon fiberreinforced polymer (CFRP), glass fiber reinforced polymer (GFRP) and thelike. As another example embodiment, the spars 102 may be made of metal,such as aluminum, or metal allow, such as aluminum alloy. In otherembodiments, the spars 102 may also be made of another suitable materialor combination of materials.

As an example embodiment, the ribs 128 may be made of a compositematerial. As an example, the ribs 128 may be made of fiber-reinforcedpolymer that includes a polymer matrix reinforced with fibers, such ascarbon fiber reinforced polymer CFRP, GFRP and the like. As anotherexample embodiment, the ribs 128 may be made of metal, such as aluminum,or metal allow, such as aluminum alloy. In other embodiments, the ribs128 may also be made of another suitable material or combination ofmaterials.

Thus, in an example embodiment, the wing box 132 (e.g., the spars 102 orthe spars 102 and the ribs 128 forming the internal frame 134 of thewing 100) may be made of metal, or metal alloy. In another exampleembodiment, the wing box 132 may be made of composite material. In yetanother example embodiment, the wing box 132 may be made of acombination of metal and composite material. The wing box 132 forms theinternal substructure of the wing 100, such as in the form of theaircraft wing 1218 (FIG. 1).

While the illustrative embodiment of the wing 100 shown in FIG. 3depicts the wing box 132 as being constructed from spars 102 and ribs128 (e.g., the wing box 132 includes interconnected spars 102 and ribs128), those skilled in the art will recognize that in other embodimentsof the wing 100, the wing box 132 may be a multi-spar design formed fromonly the plurality of spars 102 (e.g., the wing box 132 includesinterconnected spars 102).

As an example embodiment, the skins 130 (e.g., the upper skin 130 and/orthe lower skin 130 b) may be made of a composite material. As anexample, the skins 130 may be made of fiber-reinforced polymer thatincludes a polymer matrix reinforced with fibers, such as carbon fiberreinforced polymer CFRP, GFRP and the like. As another exampleembodiment, the skins 130 may be made of metal, such as aluminum, ormetal allow, such as aluminum alloy. In other embodiments, the skins 130may also be made of another suitable material or combination ofmaterials.

Thus, in an example embodiment, the disclosed wing 100 (e.g., the wingbox 132 and skins 130), such as in the form of the aircraft wing 1218(FIG. 1), may be made of composite material. In another exampleembodiment, the disclosed wing 100 may be made of metal. In yet anotherexample embodiment, the wing 100 may be made of a combination of metaland composite material.

In an example embodiment, the polymer matrix of the fiber-reinforcedpolymer, or fiber-reinforced plastic, (e.g., the resin material systemof the composite material) used to make the spars 102, the ribs 128and/or the skins 130 may be a thermoplastic resin. The presentdisclosure recognizes that the use of a thermoplastic resin may providefor advantageous embodiments because the thermoplastic resin may allowthe composite material to be heated and reformed outside of an oven orautoclave. In another example embodiment, the polymer matrix of thefiber-reinforced polymer, or fiber-reinforced plastic, used to make thespars 102, the ribs 128 and/or the skins 130 may be a thermoset resin.In yet another example, the polymer matrix of the fiber-reinforcedpolymer, or fiber-reinforced plastic, used to make the spars 102, theribs 128 and/or the skins 130 may be an epoxy resin.

Depending upon the materials used to make the spars 102 and the ribs128, the wing box 132 may be constructed according to various differentmethodologies. In the various embodiments, the spars 102 and the ribs128 are coupled together to form the wing box 132 forming the internalframe 134 of the wing 100, such as in the form of the aircraft wing 1218(FIG. 1). In an example embodiment, the spars 102 and the ribs 128 maybe connected together, for example, with mechanical fasteners, to formthe wing box 132. In another example embodiment, the spars 102 and theribs 128 may be bonded together, for example, with an adhesive, to formthe wing box 132. In another example embodiment, the spars 102 and theribs 128 may be both adhesively bonded and mechanically connectedtogether to form the wing box 132. In another example embodiment, thespars 102 and the ribs 128 may be secondary bonded together to form thewing box 132. In another example embodiment, the spars 102 and the ribs128 may be co-bonded together to form the wing box 132. In anotherexample embodiment, the spars 102 and the ribs 128 may be co-cured toform the wing box 132. In yet another example embodiment, the spars 102and the ribs 128 may be further mechanically connected together (e.g.,with fasteners) when secondary bonded, co-bonded or co-cured to form thewing box 132.

As an example, in embodiments where the spars 102 and the ribs 128 aremade of metal or a combination of metal and composite material, thespars 102 and the ribs 128 may be joined together using mechanicalfasteners, adhesives (e.g., metal bonding) or a combination ofmechanical fasteners and adhesives.

As another example, in embodiments where the spars 102 and the ribs 128are made of composite material, the spars 102 and the ribs 128 may besecondary bonded together. As used herein, secondary bonding includesthe joining together, by the process of adhesive bonding, pre-curedspars 102 and pre-cured ribs 128.

As another example, in embodiments where the spars 102 and the ribs 128are made of composite material, the spars 102 and the ribs 128 may beco-bonded together. As used herein, co-bonding includes the curingtogether of the spars 102 and the ribs 128 where one of the spars 102and the ribs 128 is fully cured and the other one of the spars 102 andthe ribs 128 is uncured.

As yet another example, in embodiments where the spars 102 and the ribs128 are made of composite material, the spars 102 and the ribs 128 maybe co-cured together. As used herein, co-curing includes the curingtogether and simultaneous bonding of the spars 102 and the ribs 128where the spars 102 and the ribs 128 are uncured.

Depending upon the materials used to make the spars 102, the ribs 128and the skins 130, the wing 100 may be constructed according to variousdifferent methodologies. In the various embodiments, the skins 130 arecoupled to wing box 132 to form the wing 100, such as in the form of theaircraft wing 1218 (FIG. 1). As an example, the skins 130 (e.g., one orboth of the upper skin 130 a and/or lower skin 130 b) are coupled to thespars 102 (e.g., one or more of the front spar 102 a, the rear spar 102b and/or any intermediate spars). In an example embodiment, the skins130 may be connected to the spars 102, for example, the mechanicalfasteners, to form the wing 100. In another example embodiment, theskins 130 may be bonded to the spars 102, for example, with an adhesive,to form the wing 100. In another example, the skins 130 may be bothadhesively bonded and mechanically connected to the spars 102 to formthe wing 100. In another example embodiment, the skins 130 and the wingbox 132 (e.g., the spars 102 and the ribs 128) may be secondary bondedtogether to form the wing 100. In another example embodiment, the skins130 and the wing box 132 may be co-bonded together to form the wing 100.In another example embodiment, the skins 130 and the wing box 132 may beco-cured to form the wing 100. In yet another example embodiment, theskins 130 and the wing box 132 may be further mechanically connectedtogether (e.g., with fasteners) when secondary bonded, co-bonded orco-cured to form the wing 100.

As an example, in embodiments where the wing box 132 (e.g., the spars102 and the ribs 128) are made of metal or a combination of metal andcomposite material and the skins 130 are made of composite material, theskins 130 and the wing box 132 may be joined together using mechanicalfasteners, adhesives or a combination of mechanical fasteners andadhesives.

As another example, in embodiments where the skins 130 and the wing box132 (e.g., the spars 102 and the ribs 128) are made of compositematerial, the skins 103 and the wing box 132 may be secondary bondedtogether. As used herein, secondary bonding includes the joiningtogether, by the process of adhesive bonding, pre-cured skins 130 and apre-cured wing box 132.

As another example, in embodiments where the skins 130 and the wing box132 are made of composite material, the skins 130 and the wing box maybe co-bonded together. As used herein, co-bonding includes the curingtogether of the skins 130 and the wing box 132 where one of the skins130 and the wing box 132 is fully cured and the other one of the skins130 and the wing box 132 is uncured. The present disclosure recognizesthat co-bonding the skins 130 and the wing box 132 may provide foradvantageous embodiments because co-bonding the skins 130 and the wingbox 132 (e.g., the skins 130 and the spars 102) may form a substantiallyunitary (e.g., one part) wing 100 and may allow for elimination of thetime consuming and expensive process of surface interface inspectionsand installation of shims to fill gaps (e.g., greater than 0.005 inch)between mating surfaces of the skins 130 and the wing box 132.

As yet another example, in embodiments where the skins 130 and the wingbox 132 are made of composite material, the skins 130 and the wing box132 may be co-cured together. As used herein, co-curing includes thecuring together and simultaneous bonding of the skins 130 and the wingbox 132 where the skins 130 and the wing box 132 are uncured. Thepresent disclosure recognizes that co-curing the skins 130 and the wingbox 132 may provide for advantageous embodiments because co-curing theskins 130 and the wing box 132 (e.g., the skins 130 and the spars 102)may form a substantially unitary (e.g., one part) wing 100 and may allowfor elimination of the time consuming and expensive process of surfaceinterface inspections and installation of shims to fill gaps (e.g.,greater than 0.005 inch) between mating surfaces of the skins 130 andthe wing box 132.

In embodiments where the skins 130 and the wing box 132 are made ofcomposite material, individual components of the wing 100 (e.g., thespars 102, the ribs 128 and/or the skins 130) or the wing 100 as a wholemay be formed according to various composite layup methodologies. As anexample, the individual components of the wing 100 or the wing 100 as awhole may be formed as a dry layup in which a plurality of sheets orplies of reinforcing fibrous material each of which beingpre-impregnated with the polymer matrix material (e.g., a pre-preg tape)is laid up, for example, in a mold, and partially or fully cured. Asanother example, the individual components of the wing 100 or the wing100 as a whole may be formed as a wet layup in which a plurality ofsheets or plies of reinforcing fibrous material is laid up, for example,in a mold, and the polymer matrix material is applied to (e.g., infusedwithin) the sheets or plies of reinforcing fibrous material andpartially or fully cured. The present disclosure recognizes that the useof the wet layup process may provide for advantageous embodimentsbecause the wet layup process may allow the individual components of thewing 100 or the wing 100 as a whole to be made at a reduced material andprocessing cost.

In the various embodiments of the wing 100 disclosed herein, such as inthe form of the aircraft wing 1218 (FIG. 1), the skins 130 are used toclose out the wing 100. As used herein, the terms “close out,” “closedout” and similar terms refer to a manufacturing methodology, process orcondition of the wing 100 in which the wing box 132 and any interiorsystems 136 are completely enclosed or sandwiched between the opposedskins 130 (e.g., the upper skin 130 a and the lower skin 130 b). Inother words, a three-dimensional structure is closed out by installing afinal part or component to completely enclose and form the structure.For example, in the case of the wing 100, five of the six sides of thewing 100 are installed, for example formed by the wing box 132 and askin 130. When the final side is installed, for example, formed by theopposed skin 130, the wing 100 is “closed out.” Examples of the interiorsystems 136 include, but are not limited to, electrical systems,hydraulic systems, fuel systems, pumps, valves, fluid tubing systems andthe like. As such, the interior systems 138 are commonly referred to asstuffed system, because the wing box 132 is filled, or stuffed, with theinterior systems 138.

Thus, once the skins 130 are coupled to and close out the wing box 132,final assembly of any interior components of the wing 100 is complete.Further, use of the disclosed fastener system 200 allows the skins 130to be fastened to the wing box 132 following close out. The presentdisclosure recognizes that using the skins 130 to close out the wing box132 may provide for advantageous embodiments because using the skins 130to close out the wing box 132 may allow for open system installation andEME protection, which eliminates the complex, expensive and laborintensive process of installation of fastener parts, installation of theinterior systems 136 and injection of EME protective sealant throughaccess holes formed in the lower outer skin.

As an example, in embodiments where the skins 130 and the wing box 132are both made of composite materials and are co-bonded or co-cured, thelower skin 130 b and the wing box 132 may be co-bonded or co-curedtogether to form a cured component (e.g., a pre-cursor composite wing).A release agent may be used between the lower skin 130 b and the wingbox 132 to enable removal of the lower skin 130 b following theco-bonding or co-curing process. The interior systems 136 are installedwithin the open wing box 132 (via open systems installation) provides bythe lack of the upper skin 130 a and/or removal of the lower skin 130 b.If temporarily removed, the lower skin 130 b is then recoupled to thewing box 132. The upper skin 130 a is then coupled to the pre-cursorcomposite wing (e.g., the wing box 132 with the recoupled lower skin 130b). As an example, the upper skin 130 a may be an uncured component inwhich the cured pre-cursor composite wing and the uncured upper skin 130a are co-bonded to form the wing 100. As another example, the upper skin130 a is a cured component in which the cured pre-cursor composite wingand the cured upper skin 130 a are secondary bonded and/or mechanicallyconnected (e.g., using fasteners) together to form the wing 100. Thisprocess may be referred to as a three-quarter co-cure.

As an example, in embodiments where the skins 130 and the wing box 132are both made of composite materials and are co-bonded or co-cured, theupper skin 130 a, the lower skin 130 b and the wing box 132 may beco-bonded or co-cured together to form a cured component (e.g., the wing100). A release agent may be used between the upper skin 130 a and thewing box 132 to enable removal of the upper skin 130 a following theco-bonding or co-curing process. Similarly, a release agent may be usedbetween the lower skin 130 b and the wing box 132 to enable removal ofthe lower skin 130 b following the co-bonding or co-curing process.After removal of one or both of the upper skin 130 a and/or the lowerskin 130 b, the interior systems 136 are installed within the open wingbox 132 (via open systems installation) due to the lack of the upperskin 130 a. The upper skin 130 a and/or the lower skin 130 b are thenrecoupled to the composite wing. This process may be referred to as afull co-cure.

FIG. 4 is a schematic illustration of a cross-sectional side elevationview of an exemplary embodiment of the disclosed fastener system 200. Inthe various embodiments of the wing 100 (FIG. 3) described herein, suchas in the form of the aircraft wing 1218 (FIG. 1), a plurality offastener systems 200 are used to further couple the skins 130 (FIG. 3)to the wing box 132 (FIG. 3). As an example, a plurality of fastenersystems 200 are used to further couple (e.g., mechanically connect) oneor more of the upper skin 130 a and/or the lower skin 130 b (FIG. 3) tothe one or more spars 102 (FIG. 3).

The fastener system 200 is a two-part system. In the illustratedembodiment, the fastener system 200 includes a nut plate 202 and afastener 204. The fastener system 200 also includes a nut 206 disposedwithin the nut plate 202.

In an exemplary embodiment, the nut plate 202 includes a body 208 and acover 210. In an exemplary embodiment, the nut plate 202 (e.g., the body208 and the cover 210) is made of metal. As a specific, non-limitingexample, the nut plate 202 is made of an anti-corrosive metal such asstainless steel, zinc-plated steel, aluminum, titanium, copper nickelalloy, copper beryllium alloy and the like. As another specific,non-limiting example, the nut plate 202 may be coated with ananti-corrosive coating, such as a barrier coating or a sacrificialcoating.

The nut plate 202 includes a central nut plate axis 218. The body 208and the cover 210 are coaxial to one another along the nut plate axis218. In an example embodiment, the body 208 and the cover 210 areseparate and discrete components that are connected together. As anexample, interfacing or joining edges of the body 208 and the cover 210may be crimped together to form the nut plate 202. In another exampleembodiment, the body 208 and the cover 210 form a unitary (one-piece)member. As an example, and as described in more detail below, the nutplate 202 may be an additively manufactured component.

In the illustrated embodiment, the body 208 includes a flange 212 and asleeve 214. The sleeve 214 includes a tubular member (e.g., a hollowcylindrical member). The flange 212 includes a circular member thatextends radially outward from the sleeve 214 and forms an exteriorshoulder 226 perpendicular to the sleeve 214. The sleeve 214 extendsaxially from the flange 212 along the nut plate axis 218.

In the illustrated embodiment, the cover 210 includes a dome 216. Thedome 216 defines a hollow interior chamber 224 (e.g., an air chamber).The dome 216 extends axially from the flange 212 along the nut plateaxis 218 opposite the sleeve 214.

In the illustrated embodiment, the fastener 204 includes a shank 220, ahead 222 disposed at an end of the shank 220 and a fastener axis 232. Inan example, at least a portion of the shank 220 includes a smoothexterior surface, for example, proximate (e.g., at or near) the head222, and at least a portion of the shank 220 includes an exteriorthread, for example, covering a portion of the shank 220 proximate theother end of the shank 220 opposite the head 222. The threaded portionof the shank 220 is configured to threadingly connect to the nut 206 inorder to fasten the fastener 204 and the nut 206 together.

In the illustrated embodiment, the nut plate 202 is configured torestrict (e.g., prevent) rotational movement of the nut 206 about thenut plate axis 218. As an example, the nut plate 202 fixes a rotationalposition of the nut 206 relative to the nut plate 202 such that the nut206 remains fixed (e.g., does not rotate about the nut plate axis 218)in response to engagement and rotational movement of the fastener 204,about the fastener axis 232, to allow the fastener 204 to be fastened(e.g., threadingly connected) to the nut 206.

In the illustrated embodiment, the sleeve 214 includes a sleeve outsidediameter D1 and a sleeve inside diameter D2. The sleeve inside diameterD2 of the sleeve 214 is larger than the shank diameter D3 of the shank220 of the fastener 204. As an example, the sleeve inside diameter D2 ofthe sleeve 214 being larger than the shank diameter D3 of the shank 220allows the fastener 204 to be inserted into and through the sleeve 214in positions where the fastener axis 232 is not coaxially aligned withthe nut plate axis 218. The difference between the sleeve insidediameter D2 of the sleeve 214 and the shank diameter D3 of the shank 220defines the positional tolerance allowance for alignment of the skinfastener hole 144 and the spar fastener hole 142. The particulardimensions of the sleeve inside diameter D2 of the sleeve 214 and theshank diameter D3 of the shank 220 of the fastener 204 may varydepending upon implementation. As a specific, non-limiting example, theshank diameter D3 of the fastener 204 may be approximately 0.003 inchand the sleeve inside diameter D2 of the sleeve 214 may be approximately0.006 inch; thus, providing a 0.003 inch radial float in any directionfor alignment of the skin fastener hole 144 and the spar fastener hole142 when coupling the skin 130 to the spar 102.

Further, in the illustrated embodiment, the nut plate 202 is alsoconfigured to allow for linear movement of the nut 206 orthogonal to thenut plate axis 218. As an example, the nut plate 202 allows the nut 206to freely move (e.g., to float) in any linear direction relative to thenut plate 202 perpendicular to the nut plate axis 218. The freeorthogonal movement of the nut 206 allows a central nut axis 234 of thenut 206 to be coaxially aligned with the fastener axis 232 and matingengagement of the fastener 204 and nut 206, when the fastener 204 ispositioned within the sleeve 214 and the fastener axis 232 is notcoaxially aligned with the nut plate axis 218.

In the illustrated embodiment, the nut 206 is disposed at leastpartially within the body 208 and at least partially within the cover210. The body 208 and the cover 210 completely enclose and seal the nut206 within the nut plate 202, for example, to protect the nut 206 andthe fastening interface from contamination, such as fuel stored withinthe fuel containment region 118 (FIG. 3). In an example embodiment, thenut 206 includes a collar 236 that extends radially outward. In thisexample embodiment, the flange 212 of the body 208 includes interiorwalls that define a nut receiving recess 230 configured to accommodateand at least partially receive the collar 236 of the nut 206. A portionof the nut 206 extending axially from the collar 236 along the nut axis234 may be disposed within the dome 216 of the cover 210.

FIG. 6 is a schematic illustration of a partial cross-sectional view ofanother embodiment of the disclosed wing 100, such as in the form of theaircraft wing 1218 (FIG. 1), and the disclosed fastener system 200. Inthe illustrated embodiment, the flange 212 forms an interior shoulder,or seat, 238 and a rim 260 at least partially defining the nut receivingrecess 230. The interior shoulder 238 of the flange 212 supports thecollar 236 of the nut 206. In an example, the nut receiving recess 230(e.g., the interior surface of the flange 212) may have an interiorgeometric shape matching an exterior geometric shape of the collar 236,such as a hexagon, in order to prevent rotation of the nut 206 withinthe flange 212. In another example, the collar 236 may include a wing orother protrusion that engages a portion of the interior surface of theflange 212 in order to prevent rotation of the nut 206 within the flange212. In other examples, the interior of the flange 212 and/or the collar236 of the nut 206 may have other features that prevent rotation of thenut 206 within the flange 212.

FIG. 5 is a schematic illustration of a partial cross-sectional view ofan exemplary embodiment of the disclosed wing 100, such as in the formof the aircraft wing 1218 (FIG. 1), and the disclosed fastener system200, for example, along lines 5-5 of FIG. 3. In the illustratedembodiment, the fastener system 200 is used to fasten the skin 130 tothe wing box 132, for example, the spars 102.

In the illustrative embodiment, each of the one or more spars 102 (e.g.,the front spar 102 a, the rear spar 102 b and/or any intermediate spars)(FIG. 3) may be a C-channel spar having a C-shaped cross section. TheC-shaped cross section of the spars 102 may vary along the length of thespars 102. Only one end (e.g., an upper end) portion of the C-channelspar is illustrated in FIG. 5. Those skilled in the art will recognizethat in other embodiments, one or more of the spars 102 may have othercross-sectional shapes, such as L-shaped spars, T-shaped spars and thelike. The spar 102 includes a web portion 138 disposed between anopposed pair of chords 140, for example, a first (e.g., upper) chord 140and an opposed second (e.g., lower) chord 140. In this exampleembodiment, the C-channel spar 102 has a unitary configuration throughits entire cross-section. The chords 140 of the spar 102 are configuredto be joined to the skins 130. As an example, the upper chord 140 isconfigured to be joined to the upper skin 130 a (FIG. 3) and the lowerchord 140 is configured to the joined to the lower skin 130 b (FIG. 3).Only one (e.g., the upper) chord 140 joined to one (e.g., the upper)skin 130 is illustrated in FIG. 5.

In the illustrative embodiment, the spar 102 includes a spar fastenerhole 142 formed (e.g., drilled or otherwise machined) through the chord140. The spar fastener hole 142 is configured to accommodate (e.g.,receive) the nut plate 202. For example, the spar fastener hole 142 isconfigured to accommodate (e.g., receive) the sleeve 214. The exteriorshoulder 226 of the flange 212 defines a flange contact surface 228configured to be placed in intimate contact with a portion of a firstsurface 150 of the spar 102, for example, proximate (e.g., at or near) aperimeter of the spar fastener hole 142, when the sleeve 214 is insertedinto the spar fastener hole 142. In this embodiment, the surface 150 ofthe spar 102 and the flange contact surface 228 define a flange-to-sparinterface 254. Similarly, the skin 130 includes a skin fastener hole 144formed therethrough. The skin fastener hole 144 is configured toaccommodate (e.g., receive) the fastener 204. The spar fastener hole 142and the skin fastener hole 144 are configured to be approximatelyaligned to accommodate installation of the fastener system 200 in orderto connect the skin 130 to the spar 102.

In other embodiments, one or more of the ribs 128 may also include oneor more rib fastener holes (not explicitly illustrated) formed (e.g.,drilled or otherwise machined) through the rib 128. The rib fastenerhole may be substantially similar to the spar fastener hole 142, asdisclosed herein, in form, structure and function. As an example, therib fastener hole is configured to accommodate (e.g., receive) the nutplate 202. The flange contact surface 228 is configured to be placed inintimate contact with a portion of a surface of the rib 128, forexample, proximate (e.g., at or near) a perimeter of the rib fastenerhole, when the sleeve 214 is inserted into the rib fastener hole.

The present disclosure recognizes that the disclosed fastening system200 may provide for advantageous embodiments because the free floatingnut 206 (e.g., freely moveable orthogonal to the nut plate axis 218)(FIG. 4) may allow for determinate, or determinant, assembly (DA) of thewing 100; thus, eliminating the time consuming, complex and costlyprocess of fixture assembly of the wing 100. Fixture assembly of thewing 100 may include a match-drilling process that requires the wing box132 and the skins 130 to be assembled in a fixture, fastener holes to bedrilled through both the spars 102 and the skins 130, the skins 130 andthe wing box 132 to be pulled apart, the skins 130 and the wing box 132to be deburred or otherwise surface finished, the skins 130 and the wingbox 132 to be reassembled, and the fasteners to be fastened. Determinateassembly is a process that allows for quicker, simpler and less costlyassembly of the wing 100 by using fastener holes formed in the skins 130and the spars 102, for example, that are pre-drilled based on a pattern,to quickly align the skins 130 and the spars 102 without the use ofadditional tooling to aid with alignment.

In an example embodiment, the spar fastener holes 142 may be pre-drilledthrough the chords 140 of the spars 102 and the skin fastener holes 144may be pre-drilled through the skins 130. One or both of the sparfastener holes 142 and the skin fastener holes 144 may be full sizeholes, for example, not needing any further drilling during constructionof the wing 100. The spar fastener holes 142 include a spar fastenerhole diameter D4 and the skin fastener holes 144 include a skin fastenerhold diameter D5. In the illustrated embodiment, the spar fastener holediameter D4 of the spar fastener holes 142 is larger than the skinfastener hole diameter D5 of the skin fastener holes 144. The sparfastener hole diameter D4 of the spar fastener holes 142 isapproximately equal to the sleeve outside diameter D1 (FIG. 4) of thesleeve 214. The skin fastener hole diameter D5 of the skin fastenerholes 144 is approximately equal to the shank diameter D3 (FIG. 4) ofthe fastener 204.

The present disclosure recognizes that the disclosed wing 100 mayprovide for advantageous embodiments because the spar fastener holediameter D4 of the spar fastener holes 142 being larger than the skinfastener hole diameter D5 of the skin fastener holes 144 may allow forthe skins 130 to be fastened to the spars 102 without coaxial alignmentof the center axes of the spar fastener holes 142 and the skin fastenerholes 144 using the disclosed fastener system 200, which may allow thefastener 204 to be fastened to the nut 206 without coaxial alignment ofthe nut plate axis 218 (FIG. 4) and the fastener axis 232 (FIG. 4). Asillustrated in FIG. 6, the nut 206 moved within the nut plate 202orthogonal to the nut plate axis 218 (FIG. 4) to coaxially align withthe fastener axis 232 (FIG. 4) and the skin fastener hole center axis148 and allow the fastener 204 to be fastened to the nut 206.

Referring to FIG. 6, in the illustrated embodiment, upon assembly of theskin 130 to the spar 102, a spar fastener hole center axis 146 of thespar fastener hole 142 and a skin fastener hole center axis 148 of theskin fastener hole 144 are not coaxially aligned. Upon installation ofthe fastener 204 through the skin fastener hole 144 and through thesleeve 214 of the nut plate 202, the fastener 204 engages the nut 206and the nut 206 moves linearly to align the nut axis 234 (FIG. 4) andthe fastener axis 232 (FIG. 4) and, also the skin fastener hole centeraxis 148 to receive the threaded end of the shank 220. In an example,the end of the fastener 204 may include a lead-in chamfer to guide thefastener 204 into the nut 206 and/or position the nut 206 relative tothe nut plate 202.

Accordingly, the disclosed fastener system 200 accounts for misalignmentof the spar fastener holes 142 and the skin fastener holes 144 that maypotentially occur using the determinate assembly process. The fastenersystem 200 enables the skin 130 to be fastened to the spar 102 with thespar fastener holes 142 and the skin fastener holes 144 not beingcoaxially aligned. Once the fastener system 200 is installed, the clampforce created by the fastener system 200 prevents any movement betweenthe skin 130 and the spar 102 due to the spar fastener hole 142 and thesleeve inside diameter D2 (FIG. 4) being greater than the shank diameterD3 (FIG. 4) of the fastener 204.

Referring now to FIGS. 5 and 6, the nut plate 202 is configured to becoupled to the spar 102, for example, to the chord 140 of the spar 102,at a fixed position with the sleeve 214 received within the sparfastener hole 142. The nut plate 202 approximately positions the nut 206relative to the spar fastener hole 142 and the skin fastener hole 144 ina suitable position for engagement with the end of the fastener 204. Asdescribed above, the nut plate 202 restricts rotational movement of thenut 206 and permits orthogonal movement of the nut 206 in order tofasten the fastener 204 to the nut 206.

The nut plate 202 may be coupled to the spar 102 by various differenttechniques. In an exemplary embodiment, the nut plate 202 is coupled tothe spar 102 using a cold expansion, or cold working, process. As anexample, prior to installation of the nut plate 202, the sleeve outsidediameter D1 of the sleeve 214 is less than the spar fastener holediameter D4 of the spar fastener hole 142. In an example operation,after the sleeve 214 is received within the spar fastener hole 142, apull gun (not shown) is operated to extend a mandrel (not shown) throughthe sleeve 214 so that a head end of the mandrel extends outwardlybeyond an outer end of the sleeve 214. The diameter of the head end ofthe mandrel plus the thickness of the sleeve 214 is approximately equalto the spar fastener hole diameter D4 of the spar fastener hole 142. Themandrel is then retracted to deform the sleeve 214 and increase thesleeve outside diameter D1 of the sleeve 214 (e.g., cold expansion) tobe approximately equal to or greater than the spar fastener holediameter D4 of the spar fastener hole 142 to hold the nut plate 202 inplace relative to the spar 102. The sleeve 214 is retained within thespar fastener hole 142 by circumferential tension about the sparfastener hole 142. In an example implementation, the nut 208 includes acounterbore that is configured to allow the mandrel to go fully throughthe sleeve 214 of the nut plate 202 in order to expand the sleeve 214without interference from the nut 208. The present disclosure recognizesthat the disclosed fastener system 200 may provide for advantageousembodiments because expansion of the sleeve 214 by the cold workingprocess may work harden the sleeve 214 and provide improved fatigue anddurability to the nut plate 202 and/or the spar 102, for example, whenthe spar 102 is made of metal.

In another example embodiment, the nut plate 202 may be adhesivelybonded to the spar 102 with the sleeve 214 positioned within the sparfastener hole 142. In yet another example embodiment, the nut plate 202may be mechanically fastened, for example, with rivets, to the spar 102with the sleeve 214 positioned within the spar fastener hole 142. Inanother example, the nut plate 202 may be integrated into the spar 102.As an example, the nut plate 202 or a portion of the nut plate 202(e.g., the body 208 of the nut plate 202) may be integrally molded intothe spar 102. The nut 206 may then be placed within the interior chamber224 formed by the integral dome 216. An insert, such as a threadedwasher, may be placed over the integral nut plate 202 to serve as therim 260 and to hold the nut 206 within the nut plate 202.

FIG. 7 is a schematic illustration of an enlarged partialcross-sectional view of another embodiment of the disclosed wing 100,such as in the form of the aircraft wing 1218 (FIG. 1), and thedisclosed fastener system 200. In an exemplary embodiment, the fastenersystem 200 is an EME-protective fastener system.

The present disclosure recognizes that the disclosed fastener system 200may provide for advantageous embodiments because use of the fastenersystem 200 to fasten the skin 130 to the spar 102 may reduce the cost,time and complexity of EME protection by eliminating the use of specialEME fasteners, EME sealant and/or other EME protection devices.

In an example embodiment, the nut plate 202 includes a dielectriccoating 240 on an interior surface 242 of the sleeve 214. The presentdisclosure recognizes that the disclosed fastener system 200 may providefor advantageous embodiments because the dielectric coating 240 providesEME protection by preventing arcing between the fastener 204 (e.g., theshank 220) and the nut plate 202 (e.g., the inner diameter surface 242of the sleeve 214). As an example, the dielectric coating includes asolid film lubricant, such as those per SAE AS5272.

In an example embodiment, the nut plate 202 includes a conductivenut-to-flange interface 244 between the nut 206 and the body 208. Theconductive nut-to-flange interface 244 establishes electrical connectionbetween the nut 206 and the nut plate 202. As an example, the nut 206includes a nut conductive contact surface 246 and the flange 212includes a flange conductive contact surface 248 that define theconductive nut-to-flange interface 244. As an example, the nutconductive contact surface 246 is defined by one or more surfaces of thecollar 236 and the flange conductive contact surface 248 is defined byone or more interior surfaces of the flange 212, for example, the rim260, forming the nut receiving recess 230. In an example, the nutconductive contact surface 246 and the flange conductive contact surface248 are both bare metal surfaces, such that the conductive nut-to-flangeinterface 244 is a metal-to-metal interface. The present disclosurerecognizes that the disclosed fastener system 200 may provide foradvantageous embodiments because the conductive nut-to-flange interface244 provides EME protection by enabling an electrical connection betweenthe nut 206 and the body 208 to allow current to flow therebetweenwithout sparking.

In an example embodiment, a lubricant 250 is applied to the threadedfastener-to-nut interface 256 between the nut 206 and the threaded endportion of the shank 220. The present disclosure recognizes that thedisclosed fastener system 200 may provide for advantageous embodimentsbecause the lubricant 250 reduces friction to lower installation forceand provides a more repeatable torque/tension relationship.

In an example embodiment, the dome 216 of the cover 210 of the nut plate202 is configured to contain a buildup of pressure resulting from anEME, such as combustion resulting from a spark. In this example, theinterior chamber 224 formed by the dome 216 of the cover 210 of the nutplate 202 includes a volume sufficient to accommodate expansion ofgases, for example, due to combustion caused by an EME. As an example,the interior chamber 224 includes an overall volume that is at leastapproximately fifty percent larger than the volume of the portion of theinterior chamber 224 occupied by the nut 206.

In the illustrated embodiment, the sleeve 214 includes a sleeve heightH. In an example embodiment, the sleeve height H is approximately equalto a spar thickness T of the spar 102 (e.g., the chord 140) (FIG. 6). Inthis embodiment, an end of the sleeve 214 is placed in intimate contactwith a portion of a first surface 152 of the skin 130, for example,proximate (e.g., at or near) a perimeter of the skin fastener hole 144,when the sleeve 214 is received within the spar fastener hole 142 andthe skin 130 is placed in an assembly position relative to the spar 102to fasten the skin 130 to the spar 102, for example, when the sparfastener hole 142 and the skin fastener hole 144 are approximatelyaligned to fasten the fastener 204 to the nut 206. In this embodiment,the surface 152 of the skin 130 and the end of the sleeve 214 define asleeve-to-skin interface 255. The present disclosure recognizes that thedisclosed fastener system 200 may provide for advantageous embodimentsbecause intimate contact between the sleeve 214 and the surface 152 ofthe skin 130 at sleeve-to-skin interface 255 may provide EME protectionby preventing the escape of high energy from within the interior chamber224 of the cover 210 of the nut plate 202, for example, due to a buildupof pressure resulting from an EME. The present disclosure recognizesthat the disclosed fastener system 200 may provide for advantageousembodiments because the intimate contact between the sleeve 214 thefirst surface 152 of the skin 130 provides EME protections bycontrolling (e.g., reducing or preventing) a gap being formed betweenthe sleeve 214 and the skin 130, which may prevent sparking between thecomponents.

In another example embodiment, the sleeve height H is less that the sparthickness T of the spar 102 (e.g., the chord 140) (FIG. 6). In thisembodiment, the end of the sleeve 214 is spaced away from the firstsurface 152 of the skin 130, when the sleeve 214 is received within thespar fastener hole 142 and the skin 130 is placed in an assemblyposition relative to the spar 102 to fasten the skin 130 to the spar102, for example, when the spar fastener hole 142 and the skin fastenerhole 144 are approximately aligned to fasten the fastener 204 to the nut206. In this embodiment, sleeve-to-skin interface 255 defines a gap (notexplicitly illustrated).

In an example embodiment, one or more interfaces between the nut plate202 and the spar 102 and/or the skin 130 include a fay seal 258. The fayseal 258 is a seal between a joint formed by opposed interfacingsurfaces. As examples, the fay seal 258 may be applied to one or more ofthe flange-to-spar interface 254, the sleeve-to-skin interface 252, asleeve to spar interface 262 and/or any other appropriate surfaceinterfaces. As an example, after proper surface preparation, a sealantis applied uniformly to one of the mating surfaces of the surfaceinterface, for example, at an approximate 10 mil thickness using anysuitable application technique. The present disclosure recognizes thatthe disclosed fastener system 200 may provide for advantageousembodiments because fay seal 258 may provide EME protection by removingopen spaces or gaps where water could be trapped, which may corrosionbetween components, and where current may cross, which may cause aspark.

In another example embodiment, the nut plate 202 is a unitary member orcomponent, for example, the body 208 and the cover 210 forms a one-piecemember. The nut 206 is disposed within the unitary nut plate 202 suchthat the collar 236 is positioned within the nut receiving recess 230 ofthe flange 212 and a portion of the nut 206 is positioned within thedome 216 of the cover 210. The present disclosure recognizes that thedisclosed fastener system 200 may provide for advantageous embodimentsbecause the unitary nut plate 202 may reduce potential sparking byminimizing joining interfaces and may provide improved EME protection,for example, to contain high energy resulting from an EME, since the nutplate 202 is sealed and there are no component interfaces or joints(e.g., crimp joints between the body 208 and the cover 210). Further,the unitary nut plate 202 provides an integral seal for the interiorchamber 224, thus eliminated the need for secondary seal caps.

In an example embodiment, the nut plate 202 is made using an additive(e.g., additive layer) manufacturing process to form the one-piecemember. In other words, the unitary nut plate 202 is an additivelymanufactured component. Additive manufacturing, also known at 3Dprinting, is consolidation process, using computer-aided manufacturing(CAM) technology, which is able to produce a functional complex part,layer-by-layer, without molds or dies. Typically, the process uses apowerful heat source, such as a laser beam or an electron beam, to melta controlled amount of metal in the form of metallic powder or wire,which is then deposited, initially, on a base plate of a work piece.Subsequent layers are then built up upon each preceding layer. In otherwords, as opposed to conventional machining processes, additivemanufacturing builds complete functional parts or, alternatively, buildsfeatures on existing components, by adding material rather than byremoving it. In this example embodiment, the nut plate 202 is builtlayer-by-layer around the nut 206.

Examples of additive manufacturing techniques include: powder bedtechnologies such as Selective Laser Melting (SLM), where metal powderis melted by a laser beam and Electron Beam Melting (EBM), where metalpowder is melted by an electron beam; blown powder technologies, alsoknown as Laser Metal Deposition or Laser cladding, where the metalpowder is blown coaxially to the laser beam, which melts the particleson a base metal to form a metallurgical bond when cooled; and SelectiveLaser Sintering, where metal powder is sintered by a laser beam

As an example, a base plate may be mounted within a powder bed and thesurface of the powder is leveled off so as to just cover the surface ofthe base plate. The laser may then be scanned over the base plate alonga path, which defines a portion of the shape of the nut plate 202.Powder is melted to this shape and solidifies to a layer of metal on thebase plate in the desired shape. The powder may then be re-leveled,slightly higher, and the process is repeated to define a continuedportion of shape of the nut plate 202, for example, the dome 216 of thecover 210 and a portion of the flange 212 defining the interior shoulder238. The nut 206 may then be placed within the interior chamber 224formed by the dome 216 such that the collar 236 is supported by theinterior shoulder 238 of the flange 212. The powder may then bere-leveled, slightly higher, and the process is repeated until theremaining portion of the shape of the nut plate has been fully formed,for example, the remaining portion of the flange 212 defining theexterior shoulder 226 and the sleeve 214.

FIG. 8 is a flow diagram illustrating an exemplary embodiment of thedisclosed method 500 for making the disclosed wing 100 (FIG. 4), such asin the form of the aircraft wing 1218 (FIG. 1).

In the illustrated embodiment, the method 500 includes the step offorming the wing box 132 (FIG. 3), as shown at block 502. As an example,the wing box 132 includes one or more spars 102 (FIG. 3) and theplurality of ribs 128 (FIG. 3) connected to the spars 102. The spars 102include the plurality of spar fastener holes 142 (FIG. 5) formed (e.g.,drilled or machined) therethrough. Each one of the spar fastener holes142 includes the spar fastener hole diameter D4 (FIG. 5).

The method 500 also includes the step of forming the skins 130 (e.g.,the upper skin 130 a and the lower skin 130 b) (FIG. 3), as shown atblock 504. As an example, the skins 130 include the plurality of skinfastener holes 144 (FIG. 5) formed therethrough. Each one of the skinfastener holes 144 includes the skin fastener hole diameter D5 (FIG. 5).The spar fastener hole diameter D4 is larger than the skin fastener holediameter D5.

The method 500 also includes the step of installing the nut plates 202(FIG. 4) of the disclosed fastener system 200 (FIG. 4) within each oneof the plurality of spar fastener holes 142 (FIG. 5), as shown at block508. As an example, each one of the nut plates 202 includes the sleeve214 (FIG. 4) configured to be received and retained within an associatedone of the spar fastener holes 142, the flange 212 (FIG. 4) extendingradially from the sleeve 214 and defining the nut receiving recess 230(FIG. 4), a dome cover 210 (FIG. 4) extending axially from the flange212 opposite the sleeve 214 and defining the interior chamber 224 (FIG.4), and the nut 206 (FIG. 4) at least partially received within the nutreceiving recess 230 and enclosed within the cover 210. The nut 206 isrestricted from rotation within the nut plate 202 about the nut plateaxis 218 (FIG. 4) and is free to move linearly within the nut plateorthogonal to the nut plate axis 218.

The method 500 also includes the step of installing one of more of theinterior systems 136 (FIG. 3) within the wing box 132, as shown at block506.

The method 500 also includes the step of sandwiching the wing box 132(FIG. 3) and enclosing the interior system 136 (FIG. 3) between theskins 130 (FIG. 3), as shown at block 510. The skin fastener holes 144(FIG. 5) are generally aligned with the spar fastener holes 142. Theskin fastener hole center axis 148 (FIG. 6) of each one of the skinfastener holes 144 is not coaxially aligned with the spar fastener holecenter axis 146 (FIG. 6) of each one of the spar fastener holes 142.

The method 500 also includes the step of installing the fasteners 204(FIG. 4), as shown at block 512. The fasteners 204 are installed througheach one of the skin fastener holes 144 (FIG. 5) and through the sleeve214 (FIG. 4) of each one of the nut plates 202 (FIG. 4) received withinassociated ones of the spar fastener holes 142.

The method 500 also includes the step of coaxially aligning the nut axis234 (FIG. 4) of the nut 206 (FIG. 4) with the skin fastener hole centeraxis 148 (FIG. 6), as shown at block 514. The nut axis 234 of the nut206 is coaxially aligned with the skin fastener hole center axis 148 ofthe associated one of the skin fastener holes 144 (FIG. 6). Coaxialalignment of the nut axis 234 and the skin fastener hole center axis 148is achieved by linearly moving the nut 206 within the nut plate 202orthogonal to the nut plate axis 218 (FIG. 4) upon engagement with thefastener 204.

The method 500 also includes the step of torqueing (e.g., fastening) thefasteners 204 (FIG. 4) to the nut 206 enclosed within the associated nutplates 202 (FIG. 4), as shown at block 516. Torqueing the fasteners 204to the nuts 206 enclosed within the nut plates 202 coupled to the spars102 fastens the skins 130 to the spars 102.

The method 500 also includes the step of providing protection from EME,as shown at block 518. As an example, protection from EME is provided bythe fastener system 200 by forming the electrically conductivenut-to-flange interface 244 (FIG. 7) between the flange conductivecontact surface 248 (FIG. 7) of the nut 206 (FIG. 7) and the nutconductive contact surface 246 (FIG. 7) of the flange 212 (FIG. 7). Asanother example, protection from EME is provided by the fastener system200 by applying the dielectric coating 240 (FIG. 7) to the insidediameter surface of the sleeve 214 (FIG. 7). As another example,protection from EME is provided by the fastener system 200 by sleeveheight H (FIG. 7) of the sleeve 214 being approximately equal to thespar thickness T (FIG. 7) of the spar 102 (FIG. 7) to form thesleeve-to-skin interface 252 (FIG. 7). As another example, protectionfrom EME is provided by the fastener system 200 by the dome 216 (FIG. 7)of the cover 210 (FIG. 7) forming the interior chamber 224 (FIG. 7)having a volume that is at least fifty percent greater than volumeoccupied by nut 206. As yet another example, protection from EME isprovided by the fastener system 200 by body 208 (FIG. 7) and the cover210 being integral to one another and forming a unitary nut plate 202;thus, sealing the interior chamber 224 and enclosing the nut 206 withinthe nut plate 202.

The method 500 also includes the step of closing out the wing 100 (FIG.3), as shown at block 520. Close out of the wing 100 is achieved byusing the skins 130 (FIG. 3) as the final close out panels of the wing100.

Accordingly, the present disclosure recognizes that the disclosed wing100 may provide for advantageous embodiments because utilizing thedisclosed fastener system 200 to fasten the skins 130 to the wing box132 may allow the skins 130 to define the final panel close out of thewing 100. The present disclosure also recognizes that the disclosedfastening system 200 may provide for advantageous embodiments becausethe nut plate 202 having the floating nut 206 may allow for determinateassembly of the wing 100, may eliminate the need for access doors orholes in the skins 130, may allow for pre-stuffed installation of theinterior systems 136 and may enable thinner wing design. Further, thedisclosed wing 100 utilizing the disclosed fastening system 200 maysimplify EME architecture while complying with EME requirements for anaircraft, for example, by eliminating seal caps, eliminating fay seals,eliminating fillet seals, eliminating surface protection such as copperfoil, dielectric tops, applique, etc. The present disclosure alsorecognizes that the disclosed wing 100 utilizing the disclosed fasteningsystem 200 may provide for advantageous embodiments by reducing thetime, complexity and cost associated with fixture assembly and matchdrilling.

Reference herein to “embodiment” means that one or more feature,structure, element, component or characteristic described in connectionwith the embodiment is included in at least one implementation of thedisclosed invention. Thus, the phrase “one embodiment,” “anotherembodiment,” and similar language throughout the present disclosure may,but do not necessarily, refer to the same embodiment. Further, thesubject matter characterizing any one embodiment may, but does notnecessarily, include the subject matter characterizing any otherembodiment.

Similarly, reference herein to “example” means that one or more feature,structure, element, component or characteristic described in connectionwith the example is included in at least one embodiment. Thus, thephrases “one example,” “another example,” and similar languagethroughout the present disclosure may, but do not necessarily, refer tothe same example. Further, the subject matter characterizing any oneexample may, but does not necessarily, include the subject mattercharacterizing any other example.

Unless otherwise indicated, the terms “first,” “second,” etc. are usedherein merely as labels, and are not intended to impose ordinal,positional, or hierarchical requirements on the items to which theseterms refer. Moreover, reference to a “second” item does not require orpreclude the existence of lower-numbered item (e.g., a “first” item)and/or a higher-numbered item (e.g., a “third” item).

As used herein, the phrase “at least one of”, when used with a list ofitems, means different combinations of one or more of the listed itemsmay be used and only one of the items in the list may be needed. Theitem may be a particular object, thing, or category. In other words, “atleast one of” means any combination of items or number of items may beused from the list, but not all of the items in the list may berequired. For example, “at least one of item A, item B, and item C” maymean item A; item A and item B; item B; item A, item B, and item C; oritem B and item C. In some cases, “at least one of item A, item B, anditem C” may mean, for example and without limitation, two of item A, oneof item B, and ten of item C; four of item B and seven of item C; orsome other suitable combination.

In FIGS. 2 and 8, referred to above, the blocks may represent operationsand/or portions thereof and lines connecting the various blocks do notimply any particular order or dependency of the operations or portionsthereof. Blocks, if any, represented by dashed lines indicatealternative operations and/or portions thereof. Dashed lines, if any,connecting the various blocks represent alternative dependencies of theoperations or portions thereof. It will be understood that not alldependencies among the various disclosed operations are necessarilyrepresented. FIGS. 2 and 8 and the accompanying disclosure describingthe operations of the disclosed methods set forth herein should not beinterpreted as necessarily determining a sequence in which theoperations are to be performed. Rather, although one illustrative orderis indicated, it is to be understood that the sequence of the operationsmay be modified when appropriate. Accordingly, modifications, additionsand/or omissions may be made to the operations illustrated and certainoperations may be performed in a different order or simultaneously.Additionally, those skilled in the art will appreciate that not alloperations described need be performed.

Although various embodiments of the disclosed apparatus, system andmethod have been shown and described, modifications may occur to thoseskilled in the art upon reading the specification. The presentapplication includes such modifications and is limited only by the scopeof the claims.

What is claimed is:
 1. A fastening system to fasten a skin to a spar ofa wing, said fastening system comprising: a threaded fastener configuredto be received through a skin fastener hole in said skin; a nut plateconfigured to be coupled within a spar fastener hole of said spargenerally aligned with said skin fastener hole, wherein said nut platecomprises a body and a cover; and a nut enclosed within said nut platebetween said body and said cover, wherein said nut is restricted fromrotation within said nut plate about a nut plate axis, and said nut isfree to move linearly within said nut plate orthogonal to said nut plateaxis.
 2. The fastening system of claim 1 wherein said body and saidcover form a unitary nut plate made by an additive manufacturingprocess.
 3. The fastening system of claim 1 wherein said body of saidnut plate comprises a sleeve configured to be received and retainedwithin said spar fastener hole and a flange extending radially from saidsleeve.
 4. The fastening system of claim 3 wherein said cover comprisesa dome extending axially from said flange opposite said sleeve.
 5. Thefastening system of claim 4 wherein said nut comprises a radial collar.6. The fastening system of claim 5 wherein said flange comprises a rimand an interior shoulder opposite said rim defining a nut receivingrecess formed within said flange.
 7. The fastening system of claim 6wherein said nut receiving recess is configured to at least partiallyreceive said collar and said interior shoulder is configured to supportsaid collar.
 8. The fastening system of claim 7 wherein said nut isconfigured to coaxially align a nut axis with a skin fastener holecenter axis of said skin fastener hole and engage said rim when saidfastener is received through said sleeve and fastened to said nut. 9.The fastening system of claim 8 wherein said sleeve comprises adielectric coating disposed on an inside diameter surface.
 10. Thefastening system of claim 9 wherein said dielectric coating comprises asolid film lubricant.
 11. The fastening system of claim 8 wherein saiddome defines a hollow interior chamber comprising a predetermined volumeof at least fifty percent more than a volume occupied by said nut. 12.The fastening system of claim 8 wherein said sleeve comprises a sleeveheight approximately equal to a spar thickness of said spar.
 13. Thefastening system of claim 8 wherein at least a portion of said rim ofsaid flange defines a flange conductive contact surface.
 14. Thefastening system of claim 13 wherein at least a portion of said collardefines a nut conductive contact surface.
 15. The fastening system ofclaim 14 wherein said flange conductive contact surface and said nutconductive contact surface define an electrically conductivenut-to-flange interface.
 16. The fastening system of claim 1 furthercomprising a lubricant between said threaded fastener and said nut. 17.The fastening system of claim 1 wherein said nut plate comprises ametallic material.
 18. The fastening system of claim 1 wherein said nutplate comprises titanium.
 19. A wing comprising the fastening system ofclaim
 1. 20. A method for making a wing, said method comprising: forminga wing box comprising interconnected spars and a plurality of sparfastener holes formed through said spars, each one of said spar fastenerholes comprising a spar fastener hole diameter; forming skins comprisinga plurality of skin fastener holes, each one of said skin fastener holescomprising a skin fastener hole diameter, wherein said spar fastenerhole diameter is larger than said skin fastener hole diameter;installing an interior system within said wing box; installing nutplates within each of said spar fastener holes, wherein each one of saidnut plates comprises: a sleeve configured to be received and retainedwithin an associated one of said spar fastener holes; a flange extendingradially from said sleeve and defining a nut receiving recess; a domecover extending axially from said flange opposite said sleeve anddefining an interior chamber; and a nut at least partially receivedwithin said nut receiving recess and enclosed within said cover, whereinsaid nut is restricted from rotation within said nut plate about a nutplate axis and is free to move linearly within said nut plate orthogonalto said nut plate axis; sandwiching said wing box and enclosing saidinterior system between said skins with said skin fastener holesgenerally aligned with said spar fastener holes, wherein a skin fastenerhole center axis of each one of said skin fastener holes is notcoaxially aligned with a spar fastener hole center axis of each one ofsaid spar fastener holes; installing fasteners through each one of saidskin fastener apertures and said sleeve of each one of said nut plates;coaxially aligning a nut axis of said nut with said skin fastener holecenter axis; fastening said fasteners to said nut of said nut plates;providing protection from electromagnetic effects; and closing out saidwing.